ULA Propellant Depot Paper (Updated)

Here’s one other interesting paper, which was presented at the SPACE 2008 conference a few months ago. As I mentioned in a previous post, I was given a copy of a few of the ULA papers before the conference, but decided to wait until after the conference to write about it. The past few months have been busy enough that it took me being home with a stomach flu today to have the time to finally write up a brief summary.

For those of you who were there for the propellant depot panel that I chaired at Space Access this year, the paper covers in more detail many of the things that Frank Zegler presented.

After an introduction where the benefits of propellant depots for the planned Constellation architecture (such as allowing the architecture to actually, you know, work…), a concept for a first-generation propellant depot was given. This concept was designed around some of the work they’ve done on their ACES stage (aka the Wide Body Centaur that I’ve written about previously), combined with some recent work on deployable sunshields.

ULA's Proposed Propellant Depot Concept

The paper hit on several of the key concepts that I’ve mentioned on this blog:

The benefit of “settled” cryogenic fluid management (CFM) instead of “zero-G” CFM. To reiterate, if you can force the propellant to assume a preselected orientation, almost all CFM tasks go from being science projects to being straightforward adaptations of terrestrial CFM techniques. Basically you want to keep liquid from going out the vent, and gas from being ingested into the transfer lines. They propose a combination of the propulsive settling that they’ve demonstrated over almost 200 Centaur flights, combined with a rotational settling technique similar to what we’ve discussed on this blog in the past. This rotational approach, and the transition to and from the axial settling to rotational settling is set to be demonstrated after the DMSP launch this next year. They’ll have almost 11klb of unused propellants to play with after delivering the primary payload, and they plan to squeeze as many experiments as possible out of that excess propellant. There’s another approach that Frank Z. and I were working on for an SBIR proposal a year ago that could potentially also work if this one doesn’t turn out.

Proper thermal design can allow for passive systems that minimize or eliminate boiloff. While you can add an active cooling system to compensate for a poor passive thermal design, it’s much better to do what you can first with a good passive design.

Almost all of the technologies for propellant depots are already developed, many of them to high TRLs. Especially if you go with settled cryo handling instead of insisting on zero-G.

If NASA opened up its lunar architecture to allow for the use of propellant depots, it would greatly expand the current demand for orbital launches. As the authors point out, even just topping-off the Earth Departure Stage’s LOX tanks would provide something like 10x the mass demand as COTS will.

They also discussed the importance of having experimental facilities for flight testing and maturation of these technologies before they’re implemented on real systems. They mention their Centaur Test Bed concept for cryogenic experiments as secondary payloads on Atlas V, but they also link to an interesting paper by Dr. Chato of Glenn Research Center about the history of suborbital and orbital flight testing of CFM technologies. I think this is one of those areas of research for which a low-cost, unmanned suborbital vehicle like we’re developing at Masten could greatly aid the development and maturation of critical spacefairing technologies.

There were a few issues I had with their presented concept that are probably worth mentioning. First, they focus on only providing LOX. While this may still be useful for NASA missions, it’s not as useful for commercial missions. Since hydrogen boils off a lot faster than LOX, not having a way to top off your LH2 tank on orbit eliminates one of the big benefits of propellant depots. Even if you don’t go with LH2 for your fuel, having both oxidizer and fuel at the depot gives you far more flexibility than just the one fluid. Of course, this disagreement is mostly just a matter of taste on my part.

My other concern is more substantial. They only briefly mention this in the paper, but the sunshields they’ve been working with use aluminized plastics. Unfortunately, the LEO environment is somewhat nasty on plastics due to atomic oxygen. In order to minimize degradation of their sunshield, as well as minimizing damage from space debris, they selected a 1300km altitude for their analysis. While this makes the sunshield work better, that altitude is not a great place for a depot operationally. First off, it’s inside the edges of the inner Van Allen belt. Once you get much higher than about 500-600km, the radiation dosage goes way up. This makes it a lot trickier on the electronics, and I don’t know if you could keep the rendezvous, docking, and transfer periods short enough in such an environment to avoid radiation damage to the crew (the point of this depot after all was for providing propellant for crewed lunar missions). Ideally, a propellant depot should be a place where you can loiter for a while in case something comes up that delays the mission. Lastly, 1300km is high enough up that there’s a significant penalty for delivery to that altitude. Especially for future potential RLVs. Now of course, a tug could relax that constraint a bit (and as I mentioned in my previous post would make operations a lot better in general). But I think the reality is that in order to close this case operationally, they really need to find a way to make the sunshield survivable at lower altitudes. For non-LEO propellant depots (L1/L2, LLO, Mars Orbit, etc) this shouldn’t be a problem, and the idea can probably be used as-is. But the concept probably needs some rethinking if they can’t get it to work at a reasonable orbital altitude.

[Update: I was able to dig up a bit of additional information about the concept. Apparently the 1300km number was somewhat arbitrary. The concept can work at more reasonable altitudes (ie 400km would probably be fine), it’s just a question of how long of a lifetime you want for the sunshield. With the petals concept as Frank explains it in the comments section, it sounds like down the road you might be able to replace the sunshield if it wears out, but depending on the lifetime, it may make more sense to just retire the module at that point and launch a new one. Basically, it sounds like a tradeoff between lower altitudes for easier access vs. more maintenance/replacement costs due to more wear. But this information more or less ends that key concern of mine with the concept.]

One last thought is that ULA is still pulling its punches on this technology. They talk about how it could help aid the existing Constellation architecture, but the reality is that once you have this technology, you could completely transform the Constellation architecture, or get rid of large chunks of it entirely. Once you have propellant depots you no longer nead super heavy lifters like Ares V. Depots allow you to store the propellants you need for long durations, so that the ESAS concerns about losing a mission if a given launch is delayed or failed are greatly reduced. Depots allow you to split propellant launches up over as many redundant launchers as you want. If you look carefully, you’ll notice that the ACES stages they mention at the end of the paper could carry quite a bit more propellant if you have a Depot to top them off than an un-topped EDS stage. And if you launch that stage dry, you can have a system that has better cryo thermal properties, much better performance overall, and it would be part of a system that was commercially useful for other markets. Once you go to a propellant depot architecture, you could launch all of the actual dry hardware from the ESAS architecture on two existing or near-term EELV Heavies, and then the rest of your launches you really don’t care about launcher reliability. Basically with a propellant depot architecture, you can keep the number of mission-critical rendezvous and docking opportunities to the same number as ESAS, while greatly increasing performance, reducing cost, and stimulating the private launch industry.

14 Responses to ULA Propellant Depot Paper (Updated)

I had a question about angular momentum. They talk about spinning up to 1E-5 or even 1E-6 G’s. However, the axis of rotation is about the axis of rotational symmetry. The moment of inertia will be greater about any perpendicular line. Over time, the system will try to achieve its lowest energy state and transfer the rotational moment into that “tumbling” mode. Is the angular momentum so small (I calculate ~940 seconds per rev or 0.064 RPM for R-2.2m at 1E-5 G.) that the ACS can keep this from happening?

I agree with your statement that 1300km altitude is a non-starter. However, many spacecraft have MLI that have survived for years, even decades, at 400-600 km altitude. (ERBS and Hubble for two instances) Kapton or other MLI should be fine, although one may have to treat the exterior surface with some thing impervious to AO. What *I* don’t like is that their demo conical sunshade is made of petals instead of continuous cones.

Comga,
Good questions. I don’t know if the ULA guys are reading this or not–I didn’t ping them to let them know I’d be reviewing their paper. I’m not positive on the spin dynamics question, but I wouldn’t be surprised if it took a little bit of RCS to keep it stable (either that or very clever mass properties design). Since they’re not doing a Zero Boil-Off tank though, there’ll be gas there, and at the rates we’re talking about, the RCS requirements should hopefully be low. Fortunately, this is the kind of concern they can probably resolve with the test they’ll be doing this spring.

As for the Kapton surviving….I’m not 100% sure. MLI works pretty well when the side exposed to the AO is aluminized. But in their case, they want the aluminized side on the inside, and the bare kapton on the outside. I don’t know for sure what altitude you have to get down to before that becomes a problem, but I wouldn’t be surprised if at 400-600km, it ended up being the major maintenance item for the depot.

About the petals… we originally had a continuous system- clearly that is optically superior since photon leaks are a bad thing. However as a practical matter the deployment of such a complex structure poses some real problems. Installation is a pain of course since it is a giant ring of folded up stuff. Once on, you have to demate the spacecraft to get it off. But more importantly the failure of one strut would almost certainly result in the failed deployment of the entire system since the structures are bound together. It also makes the validation of the system on the ground quite problematic with re-stowage a complete headache. We wanted something that was bullet-proof- not some delicate spiderweb that was literally hanging by a thread. That thing was designed to be deployed in a 1 G gravity field and has to tolerate a 2+ G flight G field once deployed. Capable of multiple deploys. We wanted something that was stiff, had predictable dynamics and we were willing to add a bit of mass to get to a high confidence state. The present design is a tribute to the talented designers at ILC who came up with some really elegant solutions to make this work.

Recognize that we are not building a space telescope- our optical efficiency doesn’t have to be perfect. We can tolerate imperfection. in fact for early missions we can tolerate a lost petal since it is exposed to the sun for only a fraction of the time. This is a first effort and we want to be able to show early success. Once we have success then we can talk about next generation optical excellence. If we have a failure on first flight we would have trouble getting a second flight! This matches our general philosophy of small steps and gradual evolution. I know it sounds unromantic and like a bunch of worry warts. It used to bug me enormously. But most of us have gone through a failure investigation and have no desire to explain to a paying customer how we just reduced his 6 year in build, most of billion dollars satellite to plasma because we made only a small mistake. But that it was “mostly successful” and that we were really trying hard. Somehow this never gets much traction. Don’t forget we once lost a mission because of a lousy set screw.
We also know that we can affect others by our success or failure. A single failure, even if caused by something unrelated, casts a pall over a technology for a long, long time. This technology is pivotal for long term cryogenic propulsion and H2/O2 cryogenic propulsion is the gateway to practical earth-moon-mars transport. We have no desire to jam the gate so to speak with a ham-handed, half baked attempt. We want others to build on our work- not be stymied by it.

I see one difficulty with the concept as outlined. If you rotate a cylindical tank about its long axis to settle the propellant, you effectively don’t have a low point in the tank to collect the propellant, since the cylinder wall is equidistant from the axis. It would be like trying to drain a tank in gravity with a perfectly flat floor, worsened by the low level of psuedogravity involved. Getting the last 10-5% into the pipe may be problematic, and that’s a lot of propellant unavailable when you’re using enough to refill an EDS stage.

I can see a number of solutions, such as using a cluster of tanks around a common axis, or ballasting the tank so the the axis of the tank is slightly offset from the axis of rotation. I would like to see these issues explicitly addressed.

Will,
I think their plan was actually to despin the tank before docking/propellant transfer and switching to axial propulsive settling. Part of the experiment they’re doing is to see if they can spin the tank up and spin it down. The RCS requirements to spin down, and then to accelerate the fluid enough to settle it is really miniscule if you can afford to give it a decent amount of time.

The missing element seems to be the actual propellant transfer. Settling seems to have been demonstrated at relatively low levels of psuedogravity. Once a RL-10 has ignited, propellant transfer occurs, but once the engine has iginited psuedogravity equals about .5 g or more.

Cryogenic propellant transfer at levels below about .5 g seems to be the the process that has still not been demonstrated. Settling the propellants with the pump silent is all well and good, but getting propellants flowing with a pump running at very low g levels is a different question.

Will,
Ah…I think I see what you’re missing. The propulsive settling isn’t done by firing the main engine. It’s done by venting some of the boiloff gases through a propulsive vent. You only need thrust levels in the low high single digit to low double digit lbf range. Even a brief pulse (a couple of seconds) at that microgee level is enough to overcome the net disturbing forces and cause the propellants to shift over to the outlet side. Occasional pulses every certain number of seconds can keep the stuff settled during the transfer. And then you just transfer by venting one tank and using the ullage pressure in the other tank to force the propellants through.

It’s actually pretty simple, and it’s basically the process they use for relighting the RL-10 each time they go to do a relight. IOW, it’s a technique they’ve demonstrated hundreds of times over the years.

Wow. It is great to get such a thorough explanation from Frank. His arguments seem rock solid. In this case, as in many, “better” would indeed be the enemy of “good enough”. Petals make great sense as does his cautious approach.

Will. with his “pseudogravity equals about 0.5 g”, is not considering your comment about using the RCS do do the axial settling. My question would be how much RCS propellant would it take to create 1E-5 to 1E-2 g for the full depot plus the EDS to drain the LOX. How does one estimate how long that would take?

Comga,
Pretty simple to estimate, really. Say the depot plus EDS stack is about 400,000lb at the start of the transfer (rough guess without looking back on the numbers). If you want 10 microgees, that equals about 4lbf. If you’re using cold gas that you’ve allowed to warm to near ambient, say you can get an Isp of about 40s (I think that’s reasonable for GOX). That would imply 1lb of GOX for every 10s. If it takes a full 30 minutes to do the transfer, that would be 1800 * 4 /40 = 180lb of GOX. Which is really nothing compared to a 400,000lb vehicle. If you needed more like 1E-4 g, you might be better of using a 40lb GOX/GH2 RCS engine. You’d get similar propellant consumption because even though you’re producing a lot more thrust you’re going to get a much higher Isp….

And the reality is you probably don’t need to provide continual thrust. You’re just trying to provide a net acceleration sufficient to keep the sump from getting uncovered. That can be done with pulses of RCS followed by coasting. Of course, if you’re doing a typical settled transfer, the receiving tank is going to be continuously venting anyway, so if you’re clever about things, there’s a good chance you can get most of your settling propellant from the actual transfill boiloff. If I’m understanding this correctly. Frank’s the genius on these things.

Jon:
I don’t think it’s as simple as you say in 9 and 11. A 30 minute transfer of, say 120 tonnes of propellant is 4 tonnes a minute. That’s roughly a firehose rate of propellant flow, and I don’t see you getting that just from ullage pressure. On the other hand, if you move the propellant the way Centaur actually does it, with a pump, that’s a lot of vibration, and you need to power the pump. Maybe you can keep the tank settled at low acceleration levels with the pump running, but that has yet to be demonstrated. Likewise with propellant going into the EDS tank at 4 tonnes a minute, is it all going to stay neatly where you want it?

I just pulled the 30 minutes number out of a hat, but using ullage pressure is definitely the way to go compared to a pump. Hold on, lets run some numbers…

So, 4 tonnes of LOX per minute gives you 67kg/s. I’m not sure what size of ducts they use for transfer, but say we’re talking a 6″ duct (pretty reasonable for something that big, if not on the small side). That gives a cross sectional area of about .018 m^2. 67kg/s of LOX is about 59L/s of flow…or about 3.3m/s.

Of course, I really don’t have any good way of knowing what sort of flow speed you want to keep things to to avoid geysering or other issues. But I could see those kind of flows being driven by ullage pressure.

Either way though, even if you do it over 2 hours instead of 30 minutes, you’re still only talking something like 650-700lb of settling propellant for 240,000lb transferred. Even if you let it take 8 hours, you’re still at barely 1% of the total transferred propellant load. And that’s if you have to leave the RCS running constantly. At 8 hours, you’re talking flow rates of 4kg/s. That’s about the LOX flow rate on Pixel’s engines.